Miniaturized green end-burning hybrid propulsion system for cubesats

ABSTRACT

A hybrid propulsion system includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The electrodes extend into the cavity. The fuel material is free of an oxidizer and is positioned in the cavity. The fuel material has a combustion surface and is exposed to the electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port provides a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber may be dominated by radiative heat transfer. Combustion of the fuel material in the combustion chamber may generate thrust of no more than 5 N at an oxidizer flow rate of no more than 5 g/s.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation to U.S. Non-provisional applicationSer. No. 16/945,473, filed on Jul. 30, 2020, and entitled “MiniaturizedGreen End-Burning Propulsion System for Cubesats,” which is incorporatedby this reference in its entirety

U.S. Non-provisional application Ser. No. 16/945,473 is acontinuation-in-part to U.S. Non-provisional application Ser. No.16/362,430, filed on Mar. 22, 2019, and entitled “Methods and Systemsfor Restartable, Hybrid-Rockets,” which is incorporated by thisreference in its entirety.

U.S. Non-provisional application Ser. No. 16/362,430 claims priority toU.S. Provisional Application No. 62/647,401, filed on Mar. 23, 2018,entitled “Methods and Systems for Green Rockets Using a Compressed-AirOxidizer,” which is herein incorporated by this reference in itsentirety. U.S. Non-provisional application Ser. No. 16/362,430 is also acontinuation-in-part application to U.S. Non-provisional applicationSer. No. 14/802,537, filed on Jul. 17, 2015, and entitled “RestartableIgnition Devices, Systems, and Methods Thereof,” which is incorporatedby this reference in its entirety. Ser. No. 14/802,537 is now U.S. Pat.No. 10,527,004, issued Jan. 7, 2020.

U.S. Non-provisional application Ser. No. 14/802,537 claims priority toU.S. Provisional Application No. 62/026,420, filed on Jul. 18, 2014, andentitled “Restartable Ignition Devices, Systems, and Methods Thereof,”which is herein incorporated by this reference in its entirety. U.S.Non-provisional application Ser. No. 14/802,537 is also acontinuation-in-part application to U.S. Non-provisional applicationSer. No. 13/953,877, filed on Jul. 30, 2013, entitled “Multiple UseHybrid Rocket Motor,” which is hereby incorporated by reference in itsentirety.

U.S. Non-provisional application Ser. No. 13/953,877 claims priority toU.S. Provisional Application Nos. 61/677,254; 61/677,266; 61/677,418;61/677,426; and 61/677,298; all filed Jul. 30, 2012, and all of whichare hereby incorporated by reference in their entirety.

TECHNICAL FIELD

The present invention relates generally to hybrid rocket systems and,more specifically, to devices, systems and methods of an ignition andcombustion portion of a hybrid rocket system, and particularly asmall-scale hybrid rocket system.

BACKGROUND

The current state of the art for hybrid rocket ignition systems islargely based on pyrotechnic ignition methods. These methods haveserious shortcomings including the inability to initiate multiplerestarts using a single device, thus, limiting the applicability of thehybrid rocket. Other shortcomings include significant physical andenvironmental hazards. For example, making rockets safer, less toxic,and less explosive comes at a significant cost. As the propellantmaterials become less volatile, they also become increasingly difficultto ignite. Combustion of hybrid propellant must be initiated by anigniter that provides sufficient heat to cause pyrolysis of the solidfuel grain at the head end of the motor, while simultaneously providingsufficient residual energy to overcome the combustion activation energyto initiate combustion. Thus, hybrid rockets have typically used large,high output pyrotechnic charges to initiate combustion. Such ignitersare capable of producing very high-enthalpy outputs, but are extremelysusceptible to hazards of electromagnetic radiation and presentsignificant operational hazards. Most importantly, such pyrotechnicigniters are designed as “one-shot” devices that do not allow multiplerestart capability.

Due to the lack of a reliable on-demand ignition system, hybrid rocketshave never been seriously considered as a viable alternative forin-space propulsion. Advancements related to 3D printable plastics asalternatives to legacy solid rocket binders like HTPB has made itpossible to manufacture a structural matrix with unique electricalbreakdown properties. This discovery has allowed for the development ofa unique on-demand ignition technology for hybrid rockets.

In the current CubeSat market, there does not exist a commercial off theshelf (COTS) propulsion system with flight heritage that has thrustlevels greater than tens of millinewtons or specific impulse (Isp)levels greater than 70 seconds. Electric propulsion systems provide veryhigh Isp but low thrust levels. The kinetic power per unit thrustsurrounding electric propulsion systems are typically in the range of10-100 W/mN³ depending on selected type. These large power requirements,in addition to the low thrust of electric propulsion, either limit themission CONOPS or extend the mission timeframe incurring additionfinancial expenditures. Cold gas propulsion systems are the opposite ofelectric systems in terms of performance with Isp levels less than 70seconds. The low Isp of cold gas systems directly results in large formfactors with poor volumetric performance.

The majority of hybrid rockets developed to date are traditionalcore-burning designs with significantly higher thrust levels thandesired for small spacecraft. These designs also require a highlength-to-diameter ratio, resulting in a form factor difficult to use inCubeSats or small satellites. Accordingly, there is a technology gap inthe current CubeSat market. The gap exists in the “high” thrust and“moderate” specific impulse category. This category is traditionallyfilled by mono- or bi-propellants.

The technology gap in the CubeSat propulsion market leads to missionlimitations including spacecraft attitude control difficulty due tocentrally located thrusters, low performing cold gas systems requiringlarge tanks, and electric propulsion systems with large powerrequirements. These limitations make several mission types extremelydifficult or, in some cases, unsupportable. Such missions includerendezvous, proximity operations, formation flying, and clustermanagement. A successful demonstration of a system in this technologygap would lead to a compact, highly efficient, and market disruptingpropulsion system that that would enable CubeSats for use in criticaland far-reaching missions.

Dedicated launches of Smallsats have increased recently, thus creatingan industry need for a cost effective and high-performance propulsionsystem. Many satellite developers continue to use multi-million-dollarpropulsion systems based on a propellant known as hydrazine. The hybridpropulsion system in development at the Space Dynamics Laboratory aimsto address this prohibitive cost requirement to allow more satellitedevelopers an opportunity to incorporate propulsion into their design.Current efforts are being made by AFRL, NASA, and ECAPS to developnon-hydrazine “green” propellants. However, the costs are nearlyidentical to conventional hydrazine systems. The cost associated withthese propulsion systems has led to many Smallsats launching with nopropulsion; often restricting the mission CONOPS. In addition to thecosts surrounding hydrazine and other “green” propellants are thehazards they pose to personnel.

Hybrid rockets typically consist of moderately benign gaseous or liquidoxidizer and an inert solid fuel. Hybrid rockets possess operationalsafety and handling advantages when compared to liquid and solidpropellent systems. The U.S. Department of Transportation concluded thatmost hybrid rocket motor designs could be safely stored and operatedwithout risk of explosion or detonation, potentially offeringsignificantly lower operating and integration cost. The inherent designsafety of hybrid rockets increases the potential for rideshareopportunities when compared to traditional monopropellant systems.

In order for hybrid rockets to fill the technology gap discussed in theprevious section there are two key disadvantages that must be addressed.The first is the difficulty surrounding motor ignition. A key reason whyhybrid rockets are considered to be safe is due to the stability oftheir fuel. This stability makes hybrid rockets difficult to ignite. Thehybrid rocket ignition source must provide sufficient heat to vaporizethe solid fuel grain at the head end of the motor while simultaneouslyproviding sufficient residual energy to overcome the activation energyof the oxidizers. Hybrid rockets typically use a pyrotechnic devicecalled a “squib” that ignites a secondary solid fuel motor to initializethe combustion of the hybrid motor. These squibs are often susceptibleto the hazards of electromagnetic radiation and can present a potentialexplosion hazard. Also, pyrotechnic ignitors can only be fired once,thus limiting their application as an in-space propulsive device.

The second disadvantage is that the internal motor ballistics of hybridcombustion produce regression rates that are typically 25-30% lower whencompared to solid fuel motors of the same thrust and impulse. To make upfor the lower regression rate, a higher oxidizer flow rate is requiredto maintain the same thrust level. This increases the systemsoxidizer-to-fuel ratio (O/F) and ultimately results in poor mass impulseperformance, erosive fuel burning, nozzle erosion, reduced motor dutycycles, potential combustion instability, and poorer overall performanceof the system. To overcome this O/F problem, hybrid rockets aretraditionally designed with cylindrical fuel ports that have longlength-to-diameter ratios. This high aspect ratio can result in poorvolumetric efficiency and thus limit a hybrid motor's application to acustomarily volume constrained small satellite.

BRIEF SUMMARY OF THE INVENTION

The inventors of the present disclosure have identified that it would beadvantageous to provide a hybrid rocket ignition system that has restartcapability that also is safe, less toxic, and less explosive than thecurrent state-of-the-art rocket systems.

Embodiments of the present invention are directed to various devices,systems and methods of providing a restartable ignition device for ahybrid rocket system. For example, in one embodiment, an ignition deviceincludes a housing and at least two electrodes. The housing includes afirst side and a second side and defines a bore with an axis extendingtherethrough between the first and second sides, the bore defining aninternal surface of the housing. The at least two electrodes extendthrough the housing to the internal surface. The at least two electrodesare configured to be spaced apart so as to provide an electricalpotential field along the internal surface between the at least twoelectrodes. Such housing is formed with and includes multiple flatlayers such that the multiple flat layers provide ridges along theinternal surface. With this arrangement, the internal surface with theridges are configured to concentrate an electrical charge upon beingsubjected to the electrical potential field.

In one embodiment, the ridges, under the electrical potential field, actas miniature electrodes to arc the electrical charge. In anotherembodiment, the internal surface includes grooves, each groove extendingbetween two adjacently extending ridges. In still another embodiment,each of the ridges are a periphery of each of the multiple flat layers.

In another embodiment, the multiple flat layers each define a planeoriented transverse relative to the axis of the bore. In anotherembodiment, the at least two electrodes define a line therebetween, theline being generally parallel with a plane defined by each of themultiple flat layers.

In another embodiment, the internal surface defines a step configurationsuch that the bore at the first side is larger than the bore at thesecond side, the step configuration exhibiting a shelf extending to ashelf notch, the shelf notch having the at least two electrodes. Inanother embodiment, the bore at the first side defines a first width andthe bore at the second side defines a second width, the first widthgreater than the second width. In yet another embodiment, the boreincludes a convergent configuration extending from the first side to thesecond side.

In another embodiment, the multiple layers are an AcrylonitrileButadiene Styrene (ABS) material. Such multiple layers may be formedwith a fused deposition modeling process or three-dimensional printing.

In accordance with another embodiment of the present invention, a methodof forming an ignition device, is provided. The method includes: forminga housing with multiple flat layers, the housing having a first side anda second side defining a bore with an axis extending through the housingand between the first and second sides such that the bore is defined byan internal surface of the housing; and positioning at least twoelectrodes to extend through the housing to the internal surface suchthat the at least two electrodes are spaced and configured to provide anelectrical potential field along the internal surface between the atleast two electrodes; wherein the forming the housing with multiple flatlayers step includes forming the internal surface to include ridges, theridges along the internal surface being configured to concentrate anelectrical charge upon being subjected to the electrical potentialfield.

In one embodiment, the method step of forming the internal surface toinclude ridges includes the step of forming multiple miniatureelectrodes configured to arc the electrical charge between the at leasttwo electrodes. In another embodiment, the method step of forming thehousing with multiple flat layers includes the step of forming themultiple flat layers in a plane oriented transverse relative to the axisof the bore. In still another embodiment, the method step of positioningthe at least two electrodes includes the step of positioning the atleast two electrodes to define a line therebetween such that the line isgenerally parallel with a plane defined by each of the multiple flatlayers.

In another embodiment, the method step of forming the housing withmultiple flat layers includes the step of forming the bore of thehousing to include a step configuration to define a shelf. In anotherembodiment, the method step of forming the housing with multiple flatlayers includes forming the bore of the housing to include a convergentconfiguration extending from the first side to the second side of thehousing.

In another embodiment, the method step of forming the housing withmultiple flat layers includes the step of forming the housing withlayers of a solid grain fuel material. In another embodiment, the methodstep of forming the housing with multiple flat layers includes formingthe housing with layers of an Acrylonitrile Butadiene Styrene (ABS)material.

In accordance with another embodiment of the present invention, a hybridrocket system is provided. The hybrid rocket system includes acontainer, an ignition portion, a solid-grain combustion portion, a postcombustion portion, and a nozzle. The container is sized to containliquid or gaseous fuel. The ignition portion includes a first side and asecond side, the ignition portion defining a bore with an axis extendingtherethrough between the first and second sides, the bore defined by aninternal surface and the bore configured to receive the fuel from thecontainer. The ignition portion includes at least two electrodesconfigured to provide an electrical potential field along the internalsurface between the at least two electrodes. The solid grain combustionportion defines a combustion chamber such that the combustion chambercorresponds with the bore of the ignition portion. The post combustionportion is coupled to the solid grain combustion portion. The nozzle iscoupled to the post combustion portion and is configured to manipulatethrust to the rocket system. The ignition portion is formed with andincludes multiple flat layers such that the multiple flat layers provideridges along the internal surface. With this arrangement, the internalsurface with the ridges are configured to concentrate an electricalcharge generally between the at least two electrodes upon beingsubjected to the electrical potential field.

In one embodiment, the multiple flat layers each define a plane orientedtransverse relative to the axis of the bore.

Another embodiment relates to a hybrid propulsion system that includes ahousing, at least two electrodes, a solid-grain fuel material, acombustion chamber, an oxidizer port, and a nozzle. The housing has afirst end and a second end and defines a cavity. The at least twoelectrodes extend into the cavity. The solid-grain fuel material is freeof an oxidizer and is positioned in the cavity. The fuel material alsohas a combustion surface and is exposed to the at least two electrodes.The combustion chamber is defined between the combustion surface and thesecond end. The oxidizer port is arranged to provide a flow of oxidizerto the combustion chamber. The nozzle is positioned at the second end.Combustion of the fuel material in the combustion chamber is dominatedby radiative heat transfer.

The hybrid propulsion system may generate no more than about 5 N ofthrust and may have oxidizer flow of no more than about 5 g/s. The fuelmaterial may include a plurality of flat layers that provide ridgesalong the combustion surface, the electrodes may be configured toconcentrate an electrical charge on the ridges, which may act asmicro-electrodes that produce localized electrical arcing thereon andignite the combustion surface of the fuel material. The fuel material atthe combustion surface may be initially consumed or removed throughcombustion of the fuel material, a newly exposed internal surface of thefuel material may have newly exposed ridges that act as newly exposedmicro-electrodes that produce localized electrical arcing thereon andmay re-ignite the newly exposed combustion surface. The fuel materialmay include a plurality of flat layers formed by an additivemanufacturing process. The oxidizer may be gaseous oxygen (GOX),hydrogen peroxide (H₂O₂), or nitrous oxide (N₂O). The fuel material mayinclude at least one of Acrylonitrile Butadiene Styrene (ABS),Polymethyl methacrylate (PMMA), Polyvinyl Chloride (PVC), and Nylon-12.The hybrid propulsion system may generate thrust in the range of about0.1 N to about 1 N. The hybrid propulsion system may use no more thanabout 1 g/s oxidizer.

In another embodiments, a hybrid propulsion system includes a housinghaving a first end and a second end and defining a cavity, at least twoelectrodes extending into the cavity, a solid-grain fuel material freeof an oxidizer and positioned in the cavity and exposed to the at leasttwo electrodes, the fuel material having a combustion surface, acombustion chamber defined between the combustion surface and the secondend, and an oxidizer port arranged to provide a flow of oxidizer to thecombustion chamber, a nozzle positioned at the second end. Combustion ofthe fuel material in the combustion chamber generates thrust of no morethan about 5 N at an oxidizer flow rate of no more than about 5 g/s.

The combustion may be dominated by radiative heat transfer. The fuelmaterial may include a plurality of flat layers that provide ridgesalong the combustion surface, and the electrodes may be configured toconcentrate an electrical charge on the ridges, which act asmicro-electrodes that produce localized electrical arcing thereon andignite the combustion surface of the fuel material. The fuel material atthe combustion surface may be initially consumed or removed throughcombustion of the fuel material, and a newly exposed internal surface ofthe fuel material may have newly exposed ridges that act as newlyexposed micro-electrodes that produce localized electrical arcingthereon and re-ignite the newly exposed combustion surface. The fuelmaterial may include a plurality of flat layers formed by an additivemanufacturing process. The oxidizer may be gaseous oxygen (GOX),hydrogen peroxide (H₂O₂), or nitrous oxide (N₂O). The fuel material mayinclude at least one of Acrylonitrile Butadiene Styrene (ABS),Polymethyl methacrylate (PMMA), Polyvinyl Chloride (PVC), and Nylon-12.The hybrid propulsion system may generate thrust in the range of about0.1 N to about 5 N. The hybrid propulsion system may use no more thanabout 5 g/s oxidizer.

In a further embodiment, a method of operating a low-thrust hybridpropulsion system includes providing a housing having first and secondends, a solid-grain fuel material positioned in the housing and having acombustion surface and being free of an oxidizer, at least twoelectrodes positioned in the housing, a combustion chamber definedbetween the combustion surface and the second end, an oxidizer port, anozzle positioned at the second end, delivering a flow of oxidizerthrough the oxidizer port and into the combustion chamber, and ignitingthe combustion surface with the at least two electrodes to generate ahot-gas, fuel-rich flow through the nozzle to generate thrust.Combustion of the fuel material in the combustion chamber is dominatedby radiative heat transfer.

The hybrid propulsion system may generate thrust in the range of about0.1 N to about 5 N. The hybrid propulsion system may generate no morethan 5 N of thrust and has oxidizer flow of no more than 5 g/s. The fuelmaterial may include a plurality of flat layers that provide ridgesalong the combustion surface, and the electrodes may be configured toconcentrate an electrical charge on the ridges, which act asmicro-electrodes that produce localized electrical arcing thereon andignite the combustion surface of the fuel material. The fuel material atthe combustion surface may be initially consumed or removed throughcombustion of the fuel material, and a newly exposed internal surface ofthe fuel material may have newly exposed ridges that act as newlyexposed micro-electrodes that produce localized electrical arcingthereon and re-ignite the newly exposed combustion surface.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The foregoing and other advantages of the invention will become apparentupon reading the following detailed description and upon reference tothe drawings in which:

FIG. 1 is a simplified side view of a hybrid rocket system, according toone embodiment of the present invention;

FIG. 2 is a perspective view of an igniter system of a hybrid rocketsystem, according to another embodiment of the present invention;

FIG. 3 is top view of the igniter system of FIG. 2, according to anotherembodiment of the present invention;

FIG. 3A is a cross-sectional view of the igniter system taken alongsection 3A of FIG. 3, according to another embodiment of the presentinvention;

FIG. 3B is a cross-sectional view of the igniter system taken alongsection 3B of FIG. 3, according to another embodiment of the presentinvention;

FIG. 4 is an enlarged view of detail 4 in FIG. 3A, depicting electrodesadjacent an internal surface, according to another embodiment of thepresent invention;

FIG. 5 is an enlarged view of detail 5 in FIG. 3B, depicting multiplelayers defined in the igniter system, according to another embodiment ofthe present invention;

FIG. 6 is an enlarged view of detail 6 in FIG. 5, depicting ridges andgrooves of the multiple layers defined in the igniter portion, accordingto another embodiment of the present invention;

FIG. 7 is a perspective view of another embodiment of an igniter system,according to the present invention;

FIG. 8 is a top view of the igniter system of FIG. 7, according toanother embodiment of the present invention; and

FIG. 8A is a cross-sectional view of the igniter system taken alongsection 8A of FIG. 8, according to another embodiment of the presentinvention.

FIG. 9 is graph showing the influence of radiative and convective heattransfer at different oxidizer mass flow levels.

FIG. 10 is a cross-sectional side view of an example hybrid rocketsystem in another embodiment of the present invention.

FIG. 11 is a cross-sectional side view of another example hybrid rocketsystem in another embodiment of the present invention.

FIG. 12 is graph showing a thrust profile of an end-burn motor inanother embodiment of the present invention.

FIGS. 13A-D are graphs showing test data from an end-burning motor inanother embodiment of the present invention.

FIGS. 14A-D are graphs showing test data from an end-burning motor inanother embodiment of the present invention.

FIG. 15 is graph showing the O/F ratio for each end burning motordiameter in another embodiment of the present invention.

FIGS. 16A-B are graphs showing thrust profiles of the core burn and endburn motors in another embodiment of the present invention.

FIGS. 17A-D are graphs showing test data from a sandwich configurationof an end-burning motor in another embodiment of the present invention.

FIG. 18 is a flow chart showing steps of an example method related tothe present invention.

DETAILED DESCRIPTION OF THE INVENTION

Referring to FIGS. 1 and 2, a simplified view of a hybrid rocket system10 or motor that includes an ignition system 12 or pre-combustionportion, according to the present invention, is provided. Referring toFIGS. 3A, 3B, and 6, in one embodiment, the ignition system 12 orpre-combustion portion may include a housing 20 formed of multiple flatlayers 18 by employing fused deposition modeling (FDM) orthree-dimensional printing. Such FDM process provides an internalsurface 14 with ridges 16 formed from the multiple flat layers 18deposited upon each other (See FIG. 6). The ignition system 12 may alsoinclude electrodes 86 and 88 spaced from each other and positionedadjacent the internal surface 14. Upon an oxidizer or oxidizer beinginjected into the system and activating an electrical potential fieldbetween the electrodes 86 and 88, the ridges 16 along the internalsurface 14 may concentrate an electrical charge which seeds combustionof the solid-grain fuel material.

As will be described herein, the unique structural characteristics ofthe material and structure of the internal surface 14 and housing 20provide an ignition system 12 that is restartable. For example, multiplere-starts have been implemented with the ignition system 12 set forthherein. The inventors have found that the only limitation to the numberof allowable restarts is the quantity of solid fuel grain materialcontained within the ignition system 12. Such ignition system 12 mayrequire small input energy and may use only non-toxic and non-explosivewith the simplicity and reliability of a monopropellant system, but withthe output enthalpy equivalent to a bi-propellant igniter. As such, therestartable ignition system 12 may have applicability to militaryaircraft, missile systems for post-stall maneuvering, emergency gasgeneration cycles, and many other applications relating to systems thatmay benefit from the restartable ignition system.

With reference to FIG. 1, the basic components of the hybrid rocketsystem 10 may include a gaseous or liquid oxidizer container 22 or tank,a combustion portion 24, and a nozzle 26. The gaseous oxidizer may benitrous oxide or gaseous oxygen or any other suitable gaseous or liquidoxidizer. The gaseous or liquid oxidizer oxidizer container 22 may bedisposed between a gas pressurization element 28 and a gas feed system30. The gas feed system 30 may feed an injector portion 32, which inturn controllably injects oxidizer into the combustion portion 24 of thehybrid rocket system 10. The combustion portion 24 of the system mayinclude multiple portions, such as, the ignition system 12 orpre-combustion portion, a main combustion portion 34, and a postcombustion portion 36. The main combustion portion 34 may be formed ofone or more solid-grain fuels, such as acrylic or hydroxyl-terminatedpolybutadiene (HTPB), or any other suitable solid fuel grain known inthe art. In one embodiment, the solid fuel grain for the main combustionportion 34 and post combustion portion 36 may be acrylonitrile butadienestyrene (ABS) or combinations of other known solid fuels. The combustionportion 24 and, more particularly, the post combustion portion 36 may becoupled to the nozzle 26 or other similar structure. The nozzle 26 mayinclude various nozzle configurations, depending upon the application ofa particular rocket system or the like. With this arrangement, theignition system 12 of the present invention may be employed with theother components of the hybrid rocket system 10 to facilitate multiplerestarts with one device, i.e., without replacing parts.

Now with reference to FIGS. 2, 3 and 3A, various views of an ignitionsystem 12 or pre-combustion portion are provided. As set forth, theignition system 12 or pre-combustion portion may be directly coupled tothe main combustion portion 34. The ignition system 12, as depicted inthe illustrated example, is directly coupled to a shortened minimalportion of the main combustion portion 34. More important to thisdescription is that the ignition system 12 or pre-combustion portion mayinclude the housing 20 and first and second electrode components 40, 42.

In one embodiment, the housing 20 may include a sleeve like structurewith various ports and notches therein and further, the sleeve likestructure may include the internal surface 14 with a step configuration.For example, the housing 20 may include a first side 44 and a secondside 46 with a bore 48 extending through and between the first andsecond sides 44, 46 of the housing 20. The second side 46 is illustratedas an interface surface between the housing 20 and main combustionportion 34. The bore 48 may define a centrally extending axis 50 along alength 52 of the housing 20. Further, the housing 20 may include anexternal surface 54 and the before mentioned internal surface 14. Theexternal surface 54 may include cylindrical shape or any anothersuitable structure.

The internal surface 14 may define the bore 48 of the housing 20, thebore 48 defining a radial component such that a cross-section of thebore 48 may be defined as generally circular or any other suitablestructure. Further, as set forth, the internal surface 14 may define astep configuration so as to include a shelf 56. In this manner, the bore48 may include a first radius 58 and a second radius 60, the firstradius 58 and the second radius 60 extending laterally from the axis 50to the internal surface 14 of the housing 20. Such first radius 58 mayextend along the length of the bore 48 from the first side 44 of thehousing 20 to the shelf 56. The second radius 60 may extend along thelength from the shelf 56 to the second side 46 of the housing 20. Withthis arrangement, the first radius 58 may be larger than the secondradius 60 such that the bore 48 exhibits a larger opening on the firstside 44 of the housing 20 than on the second side 46 of the housing 20.

With respect to FIGS. 2, 3, 3A, and 4, as set forth, the housing 20 mayinclude various ports and/or notches therein. For example, in oneembodiment, the bore 48 of the housing 20 may also include a first notch62 and a second notch 64, each defined by the internal surface 14. Thefirst and the second notches 62, 64 may be positioned on opposite sidesof the bore 48 so as to face each other. Each of the first and secondnotches 62, 64 may extend between the shelf 56 and the first side 44 ofthe housing 20 such that the shelf 56 extends further at the notch todefine a third radius 66 or a third dimension, the third radius 66 ordimension being larger than the first radius 58 and being defined fromthe axis 50 to the internal surface 14 at the first and second notches62, 64. At least one of the first and second notches 62, 64 may be sizedand configured to exhibit electrodes 86, 88 at, for example, basecorners of the at least one of the first and second notches 62, 64 andadjacent the shelf 56, discussed in further detail herein.

Further, the housing 20 may include one or more ports for the electrodecomponents. For example, the housing 20 may include a first port 68 anda second port 70. The first and second ports 68, 70 may be positionedopposite each other on the first side 44 of the housing 20. The firstport 68 may define a first port cavity 72 (shown in outline form)extending from the first port 68 to a first port outlet 74. The firstport outlet 74 may be disposed at a first base corner 76 of the firstnotch 62 on the shelf 56 and adjacent to the internal surface 14 havingthe third radius 66. Similarly, the second port 70 may extend with asecond port cavity 78 to a second port outlet 80 at a second base corner82 of the first notch 62 on the shelf 56. In this manner, the first portoutlet 74 and the second port outlet 80 of the first notch 62 may bedisposed at opposite first and second base corners 76, 82 of the firstnotch 62. A similar arrangement may be employed for the second notch 64defining first and second outlets of port cavities extending to thefirst and second ports. In this manner, the ports and cavities extendingto the first notch and/or the second notch may be sized and configuredfor positioning electrodes 86, 88 of the first and second electrodecomponents 40, 42. In another embodiment, one or both of the notches, 62or 64, or other port may include a pressure sensor configured to measurethe pressure of the oxidizer at the shelf 56.

With respect to FIGS. 3A and 4, as set forth, the ignition system 12 orpre-combustion portion includes first and second electrode components40, 42. The first and second electrode components 40, 42 may eachinclude at least a conductive electrical wire that serves as anelectrode at the end of the wire. Such electrode components may beembedded and positioned within the first and second ports 68, 70 so thatrespective first and second electrodes 86, 88 are exposed within thebore 48 and, more particularly at the first and second port outlets 74,80 defined in, for example, the first notch 62. Within the bore 48, thefirst and second electrodes 86, 88 may be spaced a distance from eachother so that, upon being electrically activated, the first and secondelectrodes 86, 88 provide a voltage potential or an electrical fieldpotential adjacent the internal surface 14 between the first and secondelectrodes 86, 88. As depicted, such distance or spacing between thefirst and second electrodes 86, 88 may be defined by the first andsecond base corners 76, 82 in, for example, the first notch 62 in thebore 48. Further, the first and second electrodes 86, 88 may be exposedat and flush with the internal surface 14 of the bore 48. In anotherembodiment, the first and second electrodes 86, 88 may protrude from theinternal surface 14 of the bore 48. Similar to that set forth above,another set of first and second electrodes 86, 88 may be positioned andspaced at the second notch 64.

As set forth, the housing 20 and bore 48 of this embodiment may includea step configuration to define the shelf 56. The shelf 56 may be sizedand configured to act as an impingement to the oxidizer or animpingement shelf to slow the oxidizer from moving down stream so as toincrease the pressure of the oxidizer at the shelf 56. The increase inpressure of the oxidizer at the shelf 56 may provide sufficient oxidizerfor a combustion reaction of a solid fuel grain material on the internalsurface 14. Suitable oxidizers may include gaseous oxygen, liquidoxygen, nitrous oxide, hydrogen peroxide, hydroxylammonium nitrate,ammonium dinitramide, or air. The oxidizer pressure increase at theimpingement shelf 56 may enable the first and second electrodes 86, 88to be minimally spaced (or minimally charged) to provide a chargeconcentration or voltage potential on the internal surface 14 of thebore 48 between the first and second electrodes 86, 88.

With respect to FIGS. 3B, 5, and 6, the housing 20 of the ignitionsystem 12 may be formed from a solid-grain fuel material. In oneembodiment, the solid-grain fuel material may be high or low densityAcrylonitrile Butadiene Styrene (ABS) or any other suitable solid-grainfuel material that holds similar electro-mechanical, combustion, andstructural properties. As set forth, the housing 20 may be formed withmultiple flat layers 18 deposited upon each other, employing the FusedDeposition Modeling (FDM) method or three-dimensional printing or anyother suitable process for layering a fuel grain. Upon employing the FDMmethod, ABS possesses a very unique electro-mechanical property suchthat additive manufacturing results in a distinctive surface structurethat is different than the surface of a monolithically fabricated (e.g.,a molded or machined) ABS structure. In particular, this surfacestructure, such as the internal surface 14 defining the bore 48, is thesurface structure that is transverse to a plane defined by any one ofthe multiple flat layers 18. Such surface structure or internal surface14 has the effect of concentrating electrical charges locally when thesurface 14 of the ABS material is subjected to an electrical potentialfield. These high-charge concentrations produce localized electricalarcing such that the ABS material breaks down at voltages significantlylower than that of a monolithically fabricated ABS structure. Describedanother way, the voltage potential created between the first and secondelectrodes 86, 88, when electrically activated, causes the uniquefeatures (the ridges 16 formed in the multiple flat layers 18 shown inFIG. 6) of the surface 14 to act as micro-electrodes which ignites thesolid-grain fuel material in the presence of an oxidizer.

In one embodiment, the multiple flat layers 18 may be deposited so thatany one of the flat layers 18 define a plane that is transverse orperpendicular with the axis 50 of the housing 20. In another embodiment,the first and second electrodes 86, 88 (see FIG. 4) may define a linetherebetween that may be generally parallel with a plane defined by eachof the multiple flat layers 18. In still another embodiment, each of theflat layers 18 may define a plane that is substantially parallel withthe axis 50 of the housing 20. In any one of these embodiments, themultiple flat layers 18, deposited upon each other, form the internalsurface 14 with ridges 16 or ridged layering. The ridges 16 or ridgedlayering may be defined by peripheral ends 90 of the multiple flatlayers 18. As set forth, the unique mechanical structure (e.g., thesurface characteristics created by the FMD layering) of the ridges 16and multiple flat layers 18, in conjunction of the material being asolid-grain fuel, such as ABS material, act as multiple micro-electrodeswhen subjected to an electrical potential field. Such unique mechanicalstructure facilitates the ignition system 12 to implement multiplerestarts. For example, even as material from the internal surface 14 isinitially consumed or removed through combustion, a newly exposedinternal surface 14 maintains similar surface characteristics or surfaceroughness that act as micro-electrodes when exposed to an electricalpotential field.

With respect to FIG. 6, an enlarged view of the multiple flat layers 18and ridges of the fuel grain material are depicted. As set forth, theinternal surface 14 defines ridges or ridged layering formed betweeneach of the multiple flat layers 18. Each of the flat layers 18 mayinclude a peak 92 with a small radius at its peripheral end such thatthe structure may also include a slope extending to the peak that may besubstantially linear or radial. Although depicted as uniform ridges 16,such ridges may not be uniform along the internal surface 14 of thehousing 20. In this manner, the internal surface 14 may exhibit a rough,coarse or scratched surface. The ridges may exhibit a nodalconfiguration or exhibit a protruding structure that may continue ordiscontinue along the peripheral end 90 of each of the multiple flatlayers 18. Likewise, the internal surface 14 may exhibit grooves 94formed between each of the multiple flat layers 18. In other words, eachgroove 94 extends between adjacently extending ridges 16. With thisarrangement, the FDM technique of forming the housing, preferably withABS material, provides for a unique electro-mechanical structure suchthat the flat layers 18 that exhibit the ridges 16 and/or grooves 94therein reacts to an electrical potential field. In this manner, thestructure and material itself act as multiple micro-electrodes, thereby,facilitating electrical breakdown to facilitate a restartable ignitionfor a hybrid rocket system.

With respect to FIGS. 7, 8, and 8A, another embodiment of an ignitionsystem 110 for a hybrid rocket system 10 (FIG. 1) is provided. Thisembodiment is similar to the previous embodiment, except this embodimentexhibits a bore 112, defined by an internal surface 114, with aconvergent or conical configuration. For example, the ignition system110 may include a housing 116 and first and second electrode components118, 119. The housing 116 may include a first side 120 and a second side122 with the bore 112 extending through and between the first and secondsides 120, 122. The bore 112 may define a centrally located axis 124extending along the length of the housing 116. The housing 116 mayinclude first and second electrode ports 126, 128 that may extend fromthe first side 120 to a convergent portion of the bore 112 so that afirst and second electrode 130, 132 may be exposed within the bore 112.The housing 116 may also include a pressure port 134 with acorresponding pressure sensor 136 so that a pressure within the bore 112may be determined upon receiving the oxidizer. Similar to that describedand depicted in FIG. 6 of the previous embodiment, the housing 116 ofthis embodiment may be formed with multiple flat layers 18 that exhibita roughened surface or ridges 16 that provide the before-discussedunique structural characteristic along the internal surface 114 of theconical bore 112. In this manner, upon the first and second electrodes130, 132 being activated to provide an electrical potential field, themultiple flat layers 18 deposited upon each other and exhibiting theridges 16 and/or grooves 94 react and concentrate a charge, thereby,acting as multiple micro-electrodes at the internal surface 114 of thebore 112.

As set forth in this embodiment, the bore 112 in the housing 116 isconvergent. The bore 112 may be sized and configured to converge so asto increase the pressure of the oxidizer as it moves downstream throughthe bore 112. The increase in pressure of the oxidizer as it movesdownstream through the bore 112 may provide sufficient oxidizer for acombustion reaction of a solid fuel-grain material on the internalsurface 114. Suitable oxidizers may include gaseous oxygen, liquidoxygen, nitrous oxide, hydrogen peroxide, hydroxylammonium nitrate,ammonium dinitramide, or air. The oxidizer pressure increase at thenarrower portion of the bore 112 may enable the first and secondelectrodes 130, 132 to be minimally spaced (or minimally charged) toprovide a charge concentration or voltage potential on the internalsurface 114 of the convergent portion of the bore 112 between the firstand second electrodes 130, 132.

Similar to previous embodiments, the multiple flat layers 18, depositedupon each other, form the internal surface 114 with ridges 16 or ridgedlayering. The unique mechanical structure (e.g., the surfacecharacteristics created by the FMD layering) of the ridges 16 andmultiple flat layers 18, in conjunction of the material being asolid-grain fuel, such as ABS material, act as multiple micro-electrodeson the internal surface 114 when subjected to an electrical potentialfield. Such unique mechanical structure facilitates the ignition system116 to implement multiple restarts. For example, even as material fromthe internal surface 114 is initially consumed or removed throughcombustion, a newly exposed internal surface 114 maintains similarsurface characteristics or surface roughness that act asmicro-electrodes when exposed to an electrical potential field fromcharged electrodes 130, 132.

Conventional hybrid rocket motors with thrust levels greater than 5 Nrely on forced convection within the boundary layer as the primary heattransfer mechanism for fuel regression. Because of convective heattransfer, the rate of fuel regression is proportional to the oxidizermass flux through the fuel port. As the fuel port burns radiallyoutwards along the motors longitudinal axis, the rate of fuel regressionresults from oxidizer diffusion into the combustion layer between theinward blowing vaporized fuel and the axially flowing oxidizer. For aconstant oxidizer mass flow, the associated mass flux and the fuelregression rate decrease with burn time. As the fuel port opens up, moresurface burn area is exposed, and the resulting oxidizer-to-fuel ratioshifts during the burn from a balance between the fuel regression ratedrop, and the increased surface burn area. Generally, the effects offorced convection tend to make hybrid motors burn from rich-to-lean overthe burn lifetime.

For Hybrid rockets with thrust levels less than 5 N, oxidizer mass flowlevels are sufficiently small that the rate of convective heat transferis significantly reduced, and radiative heat transfer dominates the fuelregression mechanism. FIG. 9 shows how hybrid rockets with loweroxidizer mass flow levels are dominated by radiative heat transfer ascompared to higher oxidizer mass flow levels being dominated byconvective heat transfer. This issue is associated with theminiaturization of hybrid rockets and has yet to be fully understood,but generally the fuel regression rate tends to grow with time. Thesesmall hybrid motors have exhibited a lean-to-rich burn behavior.

This fuel-rich tendency leads to combustion inefficiencies in the scalesnecessary to achieve useful thrust levels for small satellites (e.g.,less than 5 N). As a solution to provide constant fuel regression rateat these desired low thrust levels, an example hybrid rocket device wasredesigned to be end-burning, resulting in a constant regression rateand oxidizer-to-fuel ratio throughout the burn lifetime. This deviceprovides an end-burning hybrid thruster in the sub-Newton scale. Thedescription below relates to the design and testing results of thishybrid propulsion system at various thrust levels (e.g., 0.5 N and 1 N).

The hybrid propulsion system may use various plastics such as, forexample, ABS, PMMA, PVC, or Nylon-12 as fuel, and gaseous oxygen (GOX,)as the oxidizer. Other high-density oxidizers such as hydrogen peroxide(H₂O₂) and nitrous oxide (N₂O) may also be used. These oxidizers provideseveral advantages, including, for example, benign handling properties,simplified plumbing, and greater characteristic velocities overtraditional monopropellants.

A high voltage arc, when placed on the surface of Acrylonitrilebutadiene styrene (ABS) plastic, may result in fuel vaporization alongthe conduction path. This principle led to the development of severallab-weight thrusters varying in thrust from approximately 4-900 N. A 5 Nsystem was successfully demonstrated on a suborbital sounding rocket.During this flight demonstration, the system was fired 5 times with eachfiring lasting 5 seconds. These successful tests led an investigation ofthe possibility of filling the current CubeSat market technology gapwith this same technology by miniaturizing the ABS-based hybrid rocketsto levels less than or equal to 5 N.

An end burn configuration of a hybrid rocket system 200 (shown in FIG.10), deemed the Augmented Swirling Injection (ASI) end burn motor, wasdesigned to increase the flow of oxidizer across a fuel surface 212 of afuel gran 214 and to maintain a constant O/F ratio. The combustionchamber 216 is located at the end of the fuel grain 214, and lies inbetween the fuel grain 214 and the nozzle 218. The electrodes 220, 222run through the length of the fuel grain 214 and terminate flush withthe combustion chamber 216 surface allowing for fuel vaporization. Theoxidizer is injected at four ports 224 to induce a swirling internalflow, effectively increasing the mixing of the oxidizer with thevaporized fuel in the housing 226. A side view of the end burn design isshown in FIG. 9.

An alternative configuration, referred to as a “sandwich” configurationof a hybrid rocket system 300 (shown in FIG. 10) was designed toincrease the flow of oxidizer across the fuel surface 312 of a pair offuel grains 314 a, 314 b and to maintain a constant O/F ratio. Thecombustion chamber 316 is located between the two fuel grains 314 a, 314b, with one fuel grain 314 a bumping up against the nozzle 318 and theother fuel grain 314 b containing the electrodes 320, 322. Similar tothe ASI end-burn configuration 200, the oxidizer is injected at eightports 324 around the housing 326 to induce a swirling internal flow. Theelectrodes 320, 322 once again run through the length of the fuel grains314 a, 314 b and terminate flush with the combustion chamber 316 surfaceallowing for fuel vaporization. A side view of the end burn design isshown in FIG. 10.

One object of the hybrid rocket system 300 is to minimize the area ofthe thruster housing 326 exposed to the combustion chamber 316 with thegoal of minimizing heat loss to the housing 326. The reduction in heatloss increased combustion efficiency and fuel packing efficiency beyondwhat was demonstrated by the rocket system 200.

An issue found during testing of the system 200 was the reliability ofthe arc ignition system, including the electrodes 220, 222. While usingABS as the fuel, the arc path was repeatedly covered with the vaporizedbut unburned fuel, leading to subsequent ignition failure. Theunreliable ignition of ABS required an investigation into other plasticsto use as the fuel, including, for example, Nylon-12, PMMA (Acrylic),and PVC. These fuels were selected based on an initial analysis usingthe NASA software package Chemical Equilibrium with Applications (CEA).The CEA analysis indicated favorable combustion properties over the useof ABS. Table I shows each fuel type and the key performance metricsused to select the final fuel.

TABLE I Fuel Grain Performance Average Number of Average C* Rise timeIgnitions Fuel Type (m/s) (sec) observed ABS 895 2.18 6 Nylon-12 8602.56 17 PMMA 1022 1.86 30

Testing of ABS showed that in at least some situation it lacked theability to provide repeatable ignitions. The motor had to bedisassembled after each ignition to clean the carbon from theelectrodes. Due to the low observed variable C* and the carbon presenton the electrodes after a test, it was determined that the ABScombustion was incomplete in at least some tests.

Nylon-12 was used in several burn tests and was initially found to havebetter ignition reliability than ABS. However, the Nylon fuel wouldsoften and flow during a test, which caused ignition failure in somescenarios and, in some cases, blocked the injector port. The regressionrate in Nylon-12 was also found to have a major dependence on the bulkfuel temperature, leading to a long motor rise time as shown in FIG. 12(Thrust profile of 1 N End-Burn motor using GOX and Nylon 12).

Testing with PVC revealed a material compatibility issue between theexhaust products and the graphite nozzle, leading to nozzle erosion andfailure. A single 15 second test would produce sparks in the exhaustplume and erode the throat of a new graphite nozzle. Testing with PVCindicated that no further evaluation was required to determine its meritas a desirable fuel.

PMMA exhibited a reduced regression rate as compared to ABS, andtherefore had a reduced tendency to coat the electrodes in soot. Table Ishows that PMMA had the greatest C* value, which indicates a moreefficient combustion process when compared to ABS and Nylon-12. Becauseof the reduced deposition of soot on the electrodes during each test, agreater number of consecutive ignitions was achieved. PMMA also has ahigher melting point so it did not deform during extended durationburns. These favorable traits surrounding PMMA led to it being selectedas another desirable fuel.

Testing the arc ignition system with these plastics found that ABS wasprobably the most reliable of the plastics tested. The other plasticsusually initiated an arc path but failed soon after in subsequentignition attempts. This led to a common design of having a relativelysmall segmented portion of the primary fuel grain containing ABS as theignition system. This mitigated the above-mentioned ignition failureusing only ABS. As a result, PMMA with an ABS ignition segment was oneof the preferred motor configuration.

Two motors with different diameters of the ASI end-burning motor weretested. FIGS. 13A-D (0.5 N×4.1 cm Diameter End-Burning Motor Test Datafrom twenty-seven 5-second burns with GOX and PMMA) shows the resultsfrom over twenty burns of the 4.1 cm diameter 0.5 N motor, while FIGS.14A-D (1 N×7.5 cm Diameter End Burning IN Motor Test Data fromtwenty-seven 5-second burns with GOX and PMMA) shows the results fromover twenty burns of the 7.5 cm diameter 1 N motor. The chamberpressure, thrust, and ignition power were directly measured by the teststand, while vacuum thrust was calculated from these and the measuredoxidizer mass flow rates.

Table II shows the calculated average Isp and the extrapolated averageVacuum Isp for each ASI end-burn configuration. The Isp was calculatedover the entire duration of the burn accounting for both the rise timeand the tail-off time of the motor. The low Isp of the ASIconfigurations is a direct result of an incorrect oxidizer-to-fuelratio, which is controlled through both the geometry of the motors andthe total injected oxidizer.

TABLE II I_(sp) for 4.1 cm and 7.5 cm End Burning Motors Average AverageFuel Grain Calculated Extrapolated Diameter Ambient I_(sp) (sec) VacuumI_(sp) (sec) 4.1 cm 115 162 7.5 cm 100 133

The ignition characteristics are similar for each test as shown in FIGS.13 and 14. The chamber pressure, thrust, and Isp are not as similarbetween each burn, and may vary widely. These performance variationsrelate to the variation in the oxidizer to fuel (O/F) ratio. FIG. 15(Comparison of O/F ratios at different fuel grain diameters) shows theO/F ratio for each end burning motor diameter.

The O/F ratio tended to slightly increase with the diameter of the fuelgrain. However, the overall inconsistencies in the O/F ratio and Ispsuggest that combustion may be incomplete in these end burningconfigurations.

A comparison with a more traditional core burn configuration is shown inFIGS. 16A-B (Comparison of thrust profiles of core burn (a) and end-burn(b) motors). The core and end burning motors performed similarly to eachother at the 1 N scale. A key difference is the change in the O/F ratio.The core burn motors were extremely fuel rich, while the end burningmotors were oxidizer rich. The end burning motors successfully produce amore constant burn area and thrust level during steady state whencompared to the core burning motors. There was limited notabledifference in the Isp between the core burn and end-burn configurations.FIGS. 16A-B show a comparison of the thrust profiles of the core burnand end burn motors.

It was also discovered that a large amount of heat loss to the motorcase occurred on a burn-to-burn basis. This heat loss directly strippedaway from the combustion enthalpy, which explained much of thedifference between the performance predicted by CEA and performanceactually measured. Initial tests tracking the motor case temperatureestimated that about 20%-40% of the enthalpy released in combustion waspresent as heat in the case at the end of a test, rather than being usedto increase the chamber stagnation temperature. Two methods to try tomitigate this heat loss were developed and reached initial testing:

Method 1: Insulate the case from the combustion chamber with a hightemperature ceramic.

Method 2: Change to a “sandwich” geometry, with the combustion chamberlocated between two end-burn surfaces.

Insulating the case brought the challenge of determining a suitableinsulation material. Several ceramics were investigated with the commonproperties of withstanding high temperatures, resisting oxidizingenvironments, and having relatively low thermal conductivity. Someinitial testing was done by modifying the end burn motors to include aceramic insulator.

A “sandwich” end burn motor was also built and tested. One object ofthis approach was to minimize the area of the thruster case exposed tothe combustion chamber and increase the fuel packing efficiency. Initialtests of this motor were promising showing an increase in Isp by about23% over an end-burn configuration of the same thrust level. FIGS. 17A-D(“Sandwich” configuration showing 13×15-second burns with GOX and PMMA)show the results from the sandwich motor testing.

The heat present in the case was also reduced, averaging about 12% ofthe combustion energy over the tests performed. The thrust rise timealso improved from greater than 1 second (e.g., FIG. 13A, FIG. 14A) toabout 350 msec. These changes also improved the ignition reliability.Table III shows the average calculated Isp and the extrapolated vacuumIsp for the “Sandwich” configuration. Again, the specific impulse wascalculated over the total duration of the burn including the rise timeand tail-off time of the motor.

TABLE III “Sandwich” End Burning Motor Isp Average Average CalculatedExtrapolated Motor Ambient I_(sp) (sec) Vacuum I_(sp) (sec) “Sandwich”127 169

The propulsion systems disclosed herein take advantage of radiativeheating, which traditionally had impeded the use of hybrid rockets as apropulsion system in the 1 N class. The analysis of various benign fuelgrains paired with gaseous oxygen showed that these systems are capableof producing a vacuum Isp of greater than 150 seconds and a thrusterrise time of less than 350 msec. Other high density oxidizers such ashydrogen peroxide and nitrous oxide are expected to further increase theperformance of these end-burning hybrid motors. In at least someembodiments, the ratio between the thrust and oxidizer flow rate isfixed for any given propellant combination and ambient conditions (i.e.,vacuum optimized nozzle verses atmospheric optimized nozzle), but otheroperating conditions and losses change the ratio. For example, for ABSand GOX in vacuum, one scenario produces about 5 N using about 0.97 g/sOxygen, or about 1 N using about 0.20 g/s Oxygen. There may be similarvalues for PMMA/GOX. When using HTP or N₂O instead of GOX, there may bean increase in the oxidizer flow for a given thrust. In another example,the 1 N PMMA/GOX motor consumes about 0.6-0.7 g/s of gaseous oxygen.

FIG. 18 is a flow chart showing steps of an example method 400 inaccordance with the present disclosure. The method 400 may represent onemethod of use of one or more of the systems 200, 300 and represent amethod that is operable to generate any of the test data shown in FIGS.12-17 and/or Tables I-III disclosed herein.

The method 400 may include, at 402, providing a housing having first andsecond ends, a solid-grain fuel material positioned in the housing andhaving a combustion surface and being free of an oxidizer, at least twoelectrodes positioned in the housing, a combustion chamber definedbetween the combustion surface and the second end, an oxidizer port, anda nozzle positioned at the second end. At 404, the method 400 includesdelivering a flow of oxidizer through the oxidizer port and into thecombustion chamber. At 406, the method includes igniting the combustionsurface with the at least two electrodes to generate a hot-gas,fuel-rich flow through the nozzle to generate thrust. At 408, the methodincludes combusting of the fuel material in the combustion chamber beingdominated by radiative heat transfer.

The method 400 may further include generating thrust in the range ofabout 0.1 N to about 5 N. The method 400 may include generating no morethan about 5 N of thrust and having oxidizer flow of no more than about5 g/s. The fuel material may include a plurality of flat layers thatprovide ridges along the combustion surface, and the electrodes may beconfigured to concentrate an electrical charge on the ridges, which actas micro-electrodes that produce localized electrical arcing thereon andignite the combustion surface of the fuel material. The fuel material atthe combustion surface may initially be consumed or removed throughcombustion of the fuel material, and a newly exposed internal surface ofthe fuel material may have newly exposed ridges that act as newlyexposed micro-electrodes that produce localized electrical arcingthereon and re-ignite the newly exposed combustion surface.

While the invention may be susceptible to various modifications andalternative forms, specific embodiments have been shown by way ofexample in the drawings and have been described in detail herein.However, it should be understood that the invention is not intended tobe limited to the particular forms disclosed. Rather, the inventionincludes all modifications, equivalents, and alternatives falling withinthe spirit and scope of the invention as defined by the followingappended claims.

We claim:
 1. A hybrid propulsion system, comprising: a housing having afirst end and a second end and defining a cavity; at least twoelectrodes extending into the cavity; a solid-grain fuel material freeof oxidizer and positioned in the cavity and exposed to the at least twoelectrodes, the fuel material having a combustion surface; a combustionchamber defined between the combustion surface and the second end; anoxidizer port arranged to provide a flow of an oxidizer to thecombustion chamber; a nozzle positioned at the second end; whereincombustion of the fuel material in the combustion chamber maintainsthrust of no more than 5 N at a flow rate of the oxidizer of no morethan 5 g/s for a duration of at least between 1 and 15 seconds.
 2. Thehybrid propulsion system of claim 1, wherein the hybrid propulsionsystem has a specific impulse (Isp) greater than 100 seconds.
 3. Thehybrid propulsion system of claim 1, wherein combustion of the fuelmaterial in the combustion chamber is dominated by radiative heattransfer.
 4. The hybrid propulsion system of claim 1, wherein the fuelmaterial includes a plurality of flat layers that provide a plurality ofridges along the combustion surface, the electrodes are configured toconcentrate an electrical charge on the ridges, which act asmicro-electrodes that produce localized electrical arcing thereon andignite the combustion surface of the fuel material; and the fuelmaterial at the combustion surface is initially consumed or removedthrough combustion of the fuel material, a newly exposed internalsurface of the fuel material has newly exposed ridges that act as newlyexposed micro-electrodes that produce localized electrical arcingthereon and re-ignite the newly exposed combustion surface.
 5. Thehybrid propulsion system of claim 1, wherein the fuel material comprisesa plurality of flat layers formed by an additive manufacturing process.6. The hybrid propulsion system of claim 1, wherein the oxidizer isgaseous oxygen (GOX), hydrogen peroxide (H₂O₂), or nitrous oxide (N2O).7. The hybrid propulsion system of claim 1, wherein the fuel materialcomprises at least one of Acrylonitrile Butadiene Styrene (ABS),Polymethyl methacrylate (PMMA), Polyvinyl Chloride (PVC), and Nylon-12.8. The hybrid propulsion system of claim 1, wherein the hybridpropulsion system maintains thrust in a range of about 0.1 N to about 5N.
 9. A method of operating a low-thrust hybrid propulsion system, themethod comprising: providing a housing having first and second ends, asolid-grain fuel material positioned in the housing and having acombustion surface and being free of oxidizer, at least two electrodespositioned in the housing extending through the combustion surface, acombustion chamber defined between the combustion surface and the secondend, an oxidizer port, and a nozzle positioned at the second end;delivering a flow of oxidizer through the oxidizer port and into thecombustion chamber; igniting the combustion surface with the at leasttwo electrodes to generate a hot-gas flow through the nozzle to generatethrust; wherein combustion of the fuel material in the combustionchamber is dominated by radiative heat transfer.
 10. The method of claim9, wherein the hybrid propulsion system maintains thrust in a range ofabout 0.1 N to about 5 N for a duration of at least between 1 and 15seconds.
 11. The method of claim 9, wherein the hybrid propulsion systemmaintains combustion with no more than 5 N of thrust and has the flow ofthe oxidizer being no more than 5 g/s.
 12. The hybrid propulsion systemof claim 9, wherein the fuel material includes a plurality of flatlayers that provide a plurality of ridges along the combustion surface,the electrodes are configured to concentrate an electrical charge on theridges, which act as micro-electrodes that produce localized electricalarcing thereon and ignite the combustion surface of the fuel material,and the fuel material at the combustion surface is initially consumed orremoved through combustion of the fuel material, a newly exposedinternal surface of the fuel material has newly exposed ridges that actas newly exposed micro-electrodes that produce localized electricalarcing thereon and re-ignite the newly exposed combustion surface.